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Topic: Basic Rocket Science Q & A (Read 281651 times)

The shaped charge splits the case, the rocket motor takes care of the rest.

Ever see the Delta 241 failure? It blew up 1500 ft off the pad. It was due to a solid rocket motor failing, then triggering the destruct system. Best explaination I've heard used is a "war zone"... And yes the launch crew was int he middle of it inside a bunker when it happened.

I read a Wikipedia article on the SRB's and it quoted it using a linear shaped charge, but where exactly is the charge located, and what/how does it destroy the SRB ?

Cut it in half, cut the nozzle off etc.... ??

"The Linear Shaped Charge, installed in the systems tunnel over 70% of the motor case, cuts through the motor case to allow rapid combustion of all propellant, destroying the SRB’s. the ET, and the Orbiter"

Any other propellant combinations the readers here would like me to try?

Given the perfect impulse density which F2/NH3 offers, not to mention nice exhaust velocity, practically in the O2/H2 ballpark (see RS-68), I'd like to ask you how would the ammonia perform combined with other oxidizers, like O2, HTP, or N2O4? Would any of these combinations ever work?

strangequark, I wrote two papers on this subject which are attached. In a first stage performance is limited by propellant volume. I show in my paper that the criteria for choosing a fuel for a first stage is its impulse density Id, equal to the product of the propellant density (kg/L) and the exhuast speed (m/s). A list of propellants is given below

HTP is 98% hydrogen peroxide, RP-1 is rocket grade kerosene and C7H9 is quadricyclene or RP-X2 (an exotic hydrocarbon fuel). O2/H2 (liquid oxygen and liguid hydrogen) has the best exhaust speed, but a very poor density, which makes it a bad choice as a first stage propellant (that's why the Delta-IV is so huge). A good combination is O2/RP-1, but there are better combinations, such as HTP/RP-1 which will give 15% more performance, plus it has the advantage of being non-cryogenic at the disadvantage of being unstable in the presence of impurities.

So in answer to your specific question about O2/CH4, that performs worse than O2/RP-1 in a first stage. This means your fuel tanks will need to be about 20% larger in order to have the same performance as O2/RP-1. Against O2/H2 it performs much better. Your fuel tanks will be about 47% smaller (the actual percentage depends on the required delta-v for the first stage, larger values will decrease this amount, but even to orbital speeds, O2/CH4 will still perform better).

For an upper stage, mass is all important, so the high exhaust speed of O2/H2 means that's the best propellant to use.

By the way, for single stage to orbit, my second paper shows that any combination O2 or HTP with any other hydrocarbon fuel will outperform O2/H2. A good choice is O2/RP-1, but O2/C7H8 gives 13% more performance.

Hi Steven,

Thanks for the papers, just got the chance to really read through them. I do have a few nits to pick, however. This "density impulse" you use, it looks like it arbitrarily assumes that mass ratio scales linearly with propellant density. Tank mass seems like it roughly would (assuming: constant pressure, spherical tanks or constant height/radius cylinders), but I was curious if there's anything empirical covering the other systems as well. Or am I looking too much into it, and first stage tank mass so dwarfs the other components that a linear scaling is fair? Secondly, I was a bit thrown off by the assumption used in defining density impulse. You basically use the ln(1+x)=x assumption, but that's only valid for pretty small x, say about 0.1. So that assumption would only work for mass ratio of 1.1, at best (assuming ~450 Isp, to high-side ballpark) a first stage delta-V of just 421 m/s.

This "density impulse" you use, it looks like it arbitrarily assumes that mass ratio scales linearly with propellant density. Tank mass seems like it roughly would (assuming: constant pressure, spherical tanks or constant height/radius cylinders), but I was curious if there's anything empirical covering the other systems as well. Or am I looking too much into it, and first stage tank mass so dwarfs the other components that a linear scaling is fair?

Yes, tank mass is roughly proportional to tank volume. In the near single stage to orbit (NSTO) paper I plot engine mass (kg) to propellant flow rate (kg/s) against thrust. I found a variation from about 2 to 4 kgs/L with thrusts from 0 to 8 MN. So, there is a constant term, plus some linear term dependent on the thrust.

I believe the constant term comes about from pipe and turbo pump sizes being inversely proportional to propellant density. So for the same thrust, a low density propellant will require a bigger engine than a high density propellant, simply because the pipes and turbo pumps will be bigger. However, thrust structure mass will be proportional to thrust which is where I think the linear term comes from.

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Secondly, I was a bit thrown off by the assumption used in defining density impulse. You basically use the ln(1+x)=x assumption, but that's only valid for pretty small x, say about 0.1. So that assumption would only work for mass ratio of 1.1, at best (assuming ~450 Isp, to high-side ballpark) a first stage delta-V of just 421 m/s.

Yes, for a very close approximation, the delta-V's need to be quite small. However, if we are just interested in seeing which propellant is better as a first stage, then Figure 3 in the NSTO paper (attached below) quite clearly shows that up to about a delta-V of 3 km/s, the delta-V approximation using Id is quite valid. At higher delta-V's there is some cross over in the curves. For example, above 6 km/s O2/C7H8 performs better than HTP/C7H8. That is below dv = 6 km/s, the high Id of HTP/C7H8 allows it to perform better, while above dv = 6 km/s the better exhaust speed ve of O2/C7H8 allows it to perform better. This implies that for a traditional first stage, it is better to use HTP/C7H8. For an SSTO vehicle, O2/C7H8 is better to use, which is why I chose that propellant in this paper for my NSTO vehicle.

Updated table. We see that NH3 does not work so well with O2, N2O4 or HTP. I've also added some hydrazine N2H4 results. This gives a new record for impulse density with F2, a staggering value of 5506 Ns/L, 50% greater than O2/RP-1.

You might also note that F2 burns with water H2O quite well! F2 also burns with any hydrocarbon, literally sucking the hydrogen off and leaving a carbon soot. With H2O it leaves O2 in the exhaust! F2 also burns with HTP quite nicely, having the highest density combination, which gives it a high impulse density.

(...)Any other propellant combinations the readers here would like me to try?

Given the perfect impulse density which F2/NH3 offers, not to mention nice exhaust velocity, practically in the O2/H2 ballpark (see RS-68), I'd like to ask you how would the ammonia perform combined with other oxidizers, like O2, HTP, or N2O4? Would any of these combinations ever work?

I don't see much reason to mess with potentially explosive and/or toxic fuels or oxidizers, performance increase would not be worth it.Additional complexity of handling them securely is adding some costs, and still, can't be made 100.00% safe.

Would you like to read reports about poisonous spills from US rockets?

Would you like to put astronauts atop a rocket with 700 tons of concentrated hydrogen peroxide (as opposed to "a rocket with 700 tons of LOX")?

Thanks for reminding me about HTP. I might have to think about acquiring some for my garage project. The catalytic decomposition eases ignition, and it doesn't have any nitrogen in it to murder your Isp.

Thanks for reminding me about HTP. I might have to think about acquiring some for my garage project. The catalytic decomposition eases ignition, and it doesn't have any nitrogen in it to murder your Isp.

Be very, very careful!

HTP is catalyzed by a lot of organic sustances, including leather and human skin. Have plenty of water around to dilute any spills and a shower for any one exposed. Never work alone as the pain from exposure can prevent someone from treating themself.

IIRC, an amateur rocketry enthusiast died a couple of years ago working with HTP.

Does anyone know a good reference text for ascent trajectories? I found some old Pascal code by our very own Steven Pietrobon and some papers that didn't go into a lot of depth. I'd like to write some code to get decent payload estimates to L1 for various Jupiter variants. I have a couple of years of experience with numerical simulation, but I know next to nothing about aerodynamics.

I was always puzzled by the fact that smaller rockets can put smaller proportion of their weight into orbit. For example Falcon 1 can put 500kg, or 1.3% of launch mass, while Falcon 9 can put 10t or 3.1%. Why is that? They use same components, same manufacturing, same engines (apart from merlin<--->kestrel switch) etc.

Also Elon Musk said that they choose F1 size so that is as small as possible and capable of reaching orbit (can't remember where I read this, I'll search for it later). Does that mean that there is some lower cutoff mass for a launch vehicle?

I don't see any theoretical reason for this. For delta-v only important thing is mass fraction. I mean, what are the practical problems to make, for example, rocket with a mass of 100kg that can put like 1kg into orbit?

Does anyone know a good reference text for ascent trajectories? I found some old Pascal code by our very own Steven Pietrobon and some papers that didn't go into a lot of depth. I'd like to write some code to get decent payload estimates to L1 for various Jupiter variants. I have a couple of years of experience with numerical simulation, but I know next to nothing about aerodynamics.

Attached is a small simulation I wrote in Quick Basic. It has a very efficient numerical integration routine and the basics of aerodynamics built in. You should be able to use the aero stuff in what you are doing.

Attached is a small simulation I wrote in Quick Basic. It has a very efficient numerical integration routine and the basics of aerodynamics built in. You should be able to use the aero stuff in what you are doing.

Danny, if you or anyone else would like, I could port that code to C, C++, or Java, since some people might find it easier to read or use that way. I was thinking of doing that anyway, since I'm one of those who, despite having started with QB, finds C-like languages more readable, but it occurred to me that others might find such a translation useful...

Nikola, also for your question there are some things that every rocket has to have. A 3x rocket doesn't need a 3x flight computer. The closer one starts to that minimum function and payload rocket, the more payload one gets for a slight growth in the overall size. Look at Athena I vs Athena II and Falcon 1 vs 1e.

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